1. Field of the Invention
The present invention relates generally to gas turbine engines, and more particularly, to a nozzle assembly for a high pressure turbine of a gas turbine engine.
2. Description of Related Art
Gas turbine engines typically include an outer casing or nacelle which forms an inlet sized to provide a predetermined airflow to a core engine. The core engine generally includes a compressor for pressurizing inlet air which is discharged to a combustor where fuel is burned to provide high energy combustion gases which drive a core engine turbine or high pressure turbine. The high pressure turbine, in turn, drives the compressor.
The high pressure turbine commonly includes one or more rows or stages of a plurality of turbine blades which are circumferentially spaced apart about a turbine rotor. The high pressure turbine also includes a nozzle assembly for directing high pressure gases from the combustor onto the turbine blades at the proper angle and velocity to rotate or drive the turbine blades and rotors.
Generally, the nozzle assembly includes a plurality of nozzle vanes which extend radially between and are circumferentially spaced apart about an annular inner and outer nozzle band. The inner nozzle band provides a platform for the nozzle vane and includes a mounting flange extending radially inwardly for mounting the inner nozzle band to a nozzle support to react against the pressure loads imposed on the nozzle vanes. To provide a continuous flowpath to the platform of the nozzle vane, the inner nozzle band extends axially aft of the mounting flange to form a solid continuous lip or band "overhang".
Typically, the nozzle vanes are cooled by compressor discharged air which flows through a series of leading edge holes and gill holes or apertures on each side of the nozzle vanes. Air flowing from these holes forms a thin film of cool air over the outer surface of the nozzle vane. Internally, the nozzle vane is divided into two cavities and air flowing into the aft cavity is discharged through trailing edge slots. Moreover, the nozzle bands are cooled by impinging air on the radially inward or back side in combination with film cooling on the radially outward or gas path side.
One disadvantage of the above nozzle vane cooling scheme is that it is difficult to cool the band overhang which typically causes an area of distress. As a result, axial cracks may develop in the band overhang due to the existence of a high temperature gradient between the aft end of the band overhang and the cool mounting flange, particularly during transient operation.
One proposed solution to cool the band overhang by back side air impingement involved an impingement manifold system. An example of this is disclosed in U.S Pat. No. 4,187,054, issued Feb. 5, 1990, to inventors Landis, Jr. et al., which is assigned to the same assignee as the present invention. This manifold system allows cooling air to be introduced to the back side of the band overhang and past the mounting flange. The manifold system may be integral to the gas path side. As a result, high temperature gradient may exist between the hot gas path side and the cool back side which critically limits the life of the band overhang.
Previously , other attempts have been made to cool the band overhang, but have not proved successful or have produced poor results. One attempt was to internally cool the band overhang by serpentine cooling. An example of this is disclosed in U.S. Pat. No. 4,353,659, issued Oct. 12, 1982to inventor Hauser, which is assigned to the same present invention. Serpentine passages were cored into the nozzle band and band overhang. Air entered and flowed through the serpentine passages. Air from the serpentine passages was vented as a film upstream of a nozzle throat, the minimum passage between an adjacent pair of vanes where the gas path static pressure is the highest. However, the serpentine passages proved to be an ineffective means of conducting heat from the gas path side to the back side, resulting in high temperature gradients between the gas path side and the back side. As a result, axial cracks developed in the band overhang and cooling air leaked out through the serpentine passages rather than flowing through the complete passage. This leakage reduced both the film flow and hole backflow margin which caused the gas path side to increase in temperature, further increasing the temperature gradient.
Another attempt to cool the band overhang utilized pin-fin cooling. The back side of the band overhang was a sheet metal cover brazed onto a plurality of pins extending radially inwardly from the nozzle band. However, the sheet metal cover could not be effectively brazed onto all of the pins. As a result, heat conduction from the gas path side to the back side was localized, resulting in high temperature gradients. Additionally ,the braze material tended to fill some of the pins, resulting in hot spots.